This invention relates generally to gas turbine engines and, more particularly, to assembling rotating components of gas turbine engines.
At least some known gas turbine engines include a core engine having, in serial flow arrangement, a fan assembly and a high pressure compressor which compresses airflow entering the engine. A combustor ignites a fuel-air mixture which is then channeled through a turbine nozzle assembly towards low and high pressure turbines which each include a disk having a plurality of rotor blades that extract rotational energy from gas flow exiting the combustor. Gas turbine engines are used in different operating environments, such as, to provide propulsion for aircraft and/or to produce power in both land-based and sea-borne power systems.
During normal operations gas turbine engines may operate with high rotational speeds and relatively high temperatures. Residual stresses from a metal alloy forging process used in fabricating the turbine disks may relieve during engine operation, such that the turbine disks may undesirably expand. Moreover, such disk expansion may adversely affect rotor-to-casing clearances during operation.
To facilitate reducing occurrences of disk expansion, at least some known engine disks are spun during the manufacturing process in a near-finished condition to relieve the residual stresses. Pre-spinning of the disks has generally the same effect on relieving the residual stress as actual engine operation. Final machining, such as, of mating features and/or rabbets, for example, is accomplished after the pre-spinning process. However, the pre-spinning process may be undesirable for several reasons, such as, for example, the costs, timing and logistics associated with removing the disk from the manufacturing cycle to perform the pre-spin process. Moreover, because the high rotational speeds are needed to relieve the residual stress, a balance condition of the disk and personnel safety issues may increase the complexity of the pre-spin process.